|Publication number||US4199935 A|
|Application number||US 05/636,058|
|Publication date||Apr 29, 1980|
|Filing date||Nov 28, 1975|
|Priority date||Nov 28, 1975|
|Publication number||05636058, 636058, US 4199935 A, US 4199935A, US-A-4199935, US4199935 A, US4199935A|
|Inventors||John P. D. Hakluytt|
|Original Assignee||The Secretary Of State For Defence In Her Britannic Majesty's Government Of The United Kingdom Of Great Britain And Northern Ireland|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (6), Referenced by (7), Classifications (9)|
|External Links: USPTO, USPTO Assignment, Espacenet|
The invention relates to combustion for the combustion of fuel in an airstream and particularly, though not exclusively, relates to combustion apparatus for use in aircraft gas turbine engines.
The present invention provides a novel form of combustion apparatus wherein a high rate of heat release without excessive local heating of the apparatus can be achieved over a range of operating conditions.
According to the present invention combustion apparatus for the combustion of fuel in an airstream comprises a combustion chamber having a concave inner surface at the upstream end thereof and means for directing combustion air along the concave surface towards the middle thereof so as to form a stable vortex pattern near the concave surface, the vortical flow being in an upstream direction in the region of the periphery of the concave surface wherein combustion of the fuel takes place substantially within the vortex pattern.
Preferably, where a liquid fuel is employed, the fuel is injected at the middle of the concave surface into the combustion air flowing over the concave surface.
Embodiments of the invention will now be described, by way of example only, with reference to the drawings accompanying the Provisional Specification of which:
FIG. 1 is a sectional view of a combustion apparatus according to the invention.
FIG. 2 is a similar view of a further embodiment of the invention.
Corresponding parts in the figures are indicated by the same reference numerals.
The combustion apparatus of FIG. 1, which is the combustor of an aircraft gas turbine engine, is supported in an airstream flowing in a direction A shown in the drawing. The combustor has an annular diffuser duct 5 comprising a pair of coaxial diffuser casings 1a, 1b extending between an air inlet 6 and outlet 14 both facing upstream. The duct 5 has a return bend of annular configuration at its downstream termination. The casing 1b extends upstream of the outlet 14 to form a generally hemispherical surface over which air flows from the outlet 14 towards the middle of the surface. A series of circumferentially spaced vanes 7 located in the outlet portion of the duct 5 are inclined with respect to the longitudinal axis of the combustor and arranged so as to impart a swirling action to the air issuing from the outlet 14. A proportion of the combustion air is admitted into the combustor through an annular passage 9 defined between a tube portion forming the upstream termination of the casing 1b and a fuel injector 2 coaxially mounted within the tube portion. A flame tube 10, of smaller diameter than the diffuser duct 5, extends downstream of the duct and terminates in a nozzle 11 from which hot gas issues to a turbine (not shown). Circumferentially spaced holes 12 in the tube 10 admit dilution air into the combustor.
The combustor shown in FIG. 2 is generally similar to that shown in FIG. 1, but has a diffuser casing 1 and flame tube 10 which together form an annular member extending around the longitudinal axis B--B. The combustor has a series of circumferentially spaced fuel injectors 2.
The operation of combustors according to the invention will now be described with reference to the drawings. Air from the airstream flowing in the direction A shown in the drawings is admitted as combustion air into the inlet 6 of the diffuser duct 5 and flows through the return bend in the duct to issue from the outlet 14 onto the hemispherical surface 1b with a swirling action imparted to the flow by the vanes 7. A spray of fuel droplets (shown dotted) is directed into the air issuing from the outlet 14 by the injector 2 and a stable toroidal flow system is set up within the combustor providing a flow reversal zone in which combustion of the fuel can take place. In the embodiment of FIG. 1, a flow pattern having single toroid is set up whereas in that of FIG. 2 a pair of concentric contra rotating toroids are present. A proportion of combustion air is admitted through the passage 9 around the fuel injectors 2 and flows into the primary combustion zone in the axial direction A and assists the formation and maintenance of the toroidal flow systems. By altering the ratio of the volume flow of air through the passage 9 to that through the duct 5 the performance of the combustor can be varied.
Some improvement in air flow quality through the return bend in the duct 5 may be obtained by including further vanes in the duct at, or upstream of the bend to impart swirl to the air flowing around the bend.
The air flowing through the duct 5 cools the casing 1b. Since substantially the whole of the combustion air passes through the combustion zone the detailed design of combustion apparatus according to the invention is simplified and more effective control of the combustion process is possible.
Where the air flowing in the duct 5 has a swirl component of flow added, dilution air having a swirl component in the opposite direction may assist mixing and the development of a uniform temperature profile in the gases issuing from the nozzle 11.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US2930192 *||Dec 7, 1953||Mar 29, 1960||Gen Electric||Reverse vortex combustion chamber|
|US2974485 *||Jun 2, 1958||Mar 14, 1961||Gen Electric||Combustor for fluid fuels|
|US3352106 *||Oct 21, 1965||Nov 14, 1967||Pianko Marc||Combustion chamber with whirling slots|
|US3671171 *||Nov 27, 1970||Jun 20, 1972||Avco Corp||Annular combustors|
|FR2482E *||Title not available|
|JPS4539848B1 *||Title not available|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US4389848 *||Jan 12, 1981||Jun 28, 1983||United Technologies Corporation||Burner construction for gas turbines|
|US6564555||May 24, 2001||May 20, 2003||Allison Advanced Development Company||Apparatus for forming a combustion mixture in a gas turbine engine|
|US7410288 *||Dec 24, 1999||Aug 12, 2008||Luminis Pty. Ltd.||Fluid mixing device|
|CN103542428A *||Jul 9, 2013||Jan 29, 2014||阿尔斯通技术有限公司||Burner arrangement|
|CN103542428B *||Jul 9, 2013||Feb 10, 2016||阿尔斯通技术有限公司||炉子装置|
|EP2685171A1 *||Jul 8, 2013||Jan 15, 2014||Alstom Technology Ltd||Burner arrangement|
|EP3026347A1 *||Nov 25, 2014||Jun 1, 2016||Alstom Technology Ltd||Combustor with annular bluff body|
|U.S. Classification||60/734, 431/352|
|International Classification||F23R3/14, F23R3/54|
|Cooperative Classification||Y02T50/675, F23R3/54, F23R3/14|
|European Classification||F23R3/54, F23R3/14|